The present invention relates to components designed to operate at high temperatures. More particularly, this invention relates to methods for repair and manufacture of airfoils for gas turbine engines, and the articles made and repaired from the use of these methods.
In a gas turbine engine, compressed air is mixed with fuel in a combustor and ignited, generating a flow of hot combustion gases through one or more turbine stages that extract energy from the gas, producing output power. Each turbine stage includes a stator nozzle having vanes which direct the combustion gases against a corresponding row of turbine blades extending radially outwardly from a supporting rotor disk. The vanes and blades include airfoils having a generally concave xe2x80x9cpressurexe2x80x9d side and a generally convex xe2x80x9csuctionxe2x80x9d side, both sides extending axially between leading and trailing edges over which the combustion gases flow during operation. The vanes and blades are subject to substantial heat load, and, because the efficiency of a gas turbine engine is proportional to gas temperature, the continuous demand for efficiency improvements translates to a demand for airfoils that are capable of withstanding higher temperatures for longer service times.
Gas turbine airfoils on such components as vanes and blades are usually made of superalloys and are often cooled by means of internal cooling chambers and the addition of coatings, including thermal barrier coatings (TBC""s) and environmentally resistant coatings, to their external surfaces. The term xe2x80x9csuperalloyxe2x80x9d is usually intended to embrace iron-, cobalt-, or nickel-based alloys, which include one or more other elements including such non-limiting examples as aluminum, tungsten, molybdenum, titanium, and iron. The internal air cooling of turbine airfoils is often accomplished via a complex cooling scheme in which cooling air flows through channels within the airfoil (xe2x80x9cinternal air cooling channelsxe2x80x9d) and is then discharged through a configuration of cooling holes at the airfoil surface. Convection cooling occurs within the airfoil from heat transfer to the cooling air as it flows through the cooling channels. In addition, fine internal orifices are often provided to direct cooling air flow directly against inner surfaces of the airfoil to achieve what is referred to as impingement cooling, while film cooling is often accomplished at the airfoil surface by configuring the cooling holes to discharge the cooling air flow across the airfoil surface so that the surface is protected from direct contact with the surrounding hot gases within the engine. TBC""s comprise at least a layer of thermally insulating ceramic and often include one or more layers of metal-based, oxidation-resistant materials (xe2x80x9cenvironmentally resistant coatingsxe2x80x9d) underlying the insulating ceramic for enhanced protection of the airfoil. Environmentally resistant coatings are also frequently used without a TBC topcoat. Technologies such as coatings and internal air cooling have effectively enhanced the performance of turbine airfoils, but material degradation problems persist in turbine airfoils due to locally aggressive conditions in areas such as airfoil leading edges and trailing edges.
A considerable amount of cooling air is often required to sufficiently lower the surface temperature of an airfoil. However, the casting process and the cores required to form the cooling channels limit the complexity of the cooling scheme that can be formed within an airfoil at leading and trailing edges of vanes and blades. The resulting restrictions in cooling airflow often promote higher local temperatures in these areas relative to those existing in other locations on a given airfoil. In typical jet engines, for example, bulk average airfoil temperatures range between about 900xc2x0 C. to about 1000xc2x0 C., while airfoil leading and trailing edge surfaces often reach about 1100xc2x0 C. or more. Maximum surface temperatures are expected in future applications to be over about 1300xc2x0 C. Of particular concern is the combination of stress with temperature, because metals, including alloys used to make gas turbine airfoils, tend to become weaker, or more easily deformed, as temperatures increase. Thus, while stress of a certain level operating on a cooler section of an airfoil may have little effect on performance, the same stress level may be beyond the performance capability of the material at hotter locations as described above. At such elevated temperatures, materials are more susceptible to damage due to a number of phenomena, including diffusion-controlled deformation (xe2x80x9ccreepxe2x80x9d), cyclic loading and unloading (xe2x80x9cfatiguexe2x80x9d), chemical attack by the hot gas flow (xe2x80x9coxidationxe2x80x9d), wear from the impact of particles entrained in the gas flow (xe2x80x9cerosionxe2x80x9d), and others.
Damage to airfoils, particularly at edges, leads to degradation of turbine efficiency. As airfoils are deformed, oxidized, or worn away, gaps between components become excessively wide, allowing gas to leak through the turbine stages without the flow of the gas being converted into mechanical energy. When efficiency drops below specified levels, the turbine must be removed from service for overhaul and refurbishment. A significant portion of this refurbishment process is directed at the repair of the specific areas of airfoils. In one repair embodiment, for example, crack-filling processes based on brazing techniques are used to repair localized damage on turbine vanes.
In current practice, the original edge material is made of the same material as the rest of the original blade, often a superalloy based on nickel or cobalt. Because this material was selected to balance the design requirements of the entire blade, it is generally not optimized to meet the special local requirements demanded by conditions at the airfoil leading or trailing edges. The performance of alloys commonly used for repair is comparable or inferior to that of the material of the original component, depending upon the microstructure, defect density, and chemistry of the repair material. For example, many turbine airfoils are made using alloys that have been directionally solidified. The directional solidification process manipulates the orientation of metal crystals, or grains, as the alloy is solidified from the molten state, lining the grains up in one selected primary direction. The resultant alloy has enhanced resistance to creep and fatigue during service when compared to conventionally processed materials. Advanced applications employ alloys made of a single crystal for even further improvements in high-temperature creep and fatigue behavior. However, when these components are repaired by conventional processes, using build-up of weld filler material, the resulting microstructure of the repair is typical of welded material, not directionally solidified or single-crystalline. Other repair methods, such as applying powder mixtures wherein one powder melts and densifies the repaired area during heat treatment, results in microstructures that differ from that of the parent alloy. Such microstructures, present in a conventional airfoil material such as a superalloy, may cause the airfoil to require excessively frequent repairs in advanced designs that rely on the benefits of directional solidification or single crystal processing to maintain performance.
Materials are characterized by several properties to aid in determining their suitability for use in demanding applications such as gas turbine airfoils. The term xe2x80x9ccreep lifexe2x80x9d is used in the art to refer to the length of time until a standard specimen of material extends to a preset length or fractures when subjected to a given stress level at a given temperature. Similarly, the term xe2x80x9cfatigue lifexe2x80x9d is used in the art to describe the length of time until a standard specimen fractures when subjected to a given set of fatigue parameters, including minimum and maximum stress levels, frequency of loading/unloading cycle, and others, at a given temperature. The term xe2x80x9coxidation resistancexe2x80x9d is used in the art to refer to the amount of damage sustained by a material when exposed to oxidizing environments, such as, for example, high temperature gases containing oxygen. Oxidation resistance is generally measured as the rate at which the weight of a specimen changes per unit surface area during exposure at a given temperature. In many cases, the weight change is measured to be a net loss in weight, as metal is converted to oxide that later detaches and falls away from the surface. In other cases, a specimen may gain weight if the oxide tends to adhere to the specimen, or if the oxide forms within the specimen, underneath the surface, a condition called xe2x80x9cinternal oxidation.xe2x80x9d A material is said to have xe2x80x9chigherxe2x80x9d or xe2x80x9cgreaterxe2x80x9d oxidation resistance than another if the material""s rate of weight change per unit surface area is closer to zero than that of the other material for exposure to the same environment and temperature.
Materials particularly noted for high creep life include oxide dispersion strengthened (ODS) materials and directionally solidified eutectic (DSE) alloys. Several materials from these classes have creep lives about three times those measured for conventional superalloys. ODS materials use mechanical techniques during processing to evenly distribute hard oxide particles of sizes less than about 0.1 micron within a metallic matrix, with the particles serving as a reinforcing phase to strengthen the material. DSE alloys are characterized by carefully controlled chemistry and processing, which produce a unique microstructure comprising the inherent fibrous or, in some cases, lamellar structure of the eutectic phase, with the fibers or lamellae aligned along a desired axis of the cast part in a manner analogous to a fiber-reinforced composite. DSE materials are also notable for excellent fatigue life, with certain alloys having about three times the creep lives measured for conventional superalloys. The careful processing controls needed to produce ODS and DSE alloys cause these materials to be prohibitively expensive.
The so-called xe2x80x9cplatinum groupxe2x80x9d of metal elements comprises rhodium (Rh), osmium (Os), platinum (Pt), iridium (Ir), ruthenium (Ru), palladium (Pd), and rhenium (Re)elements noted for high chemical resistance. Several elements from this group are noteworthy as examples of materials with substantially higher oxidation resistance relative to current airfoil materials. Some platinum group metals and several alloys based on platinum group metals possess excellent resistance to oxidation at temperatures exceeding the capabilities of many Ni-based superalloys. The class of materials referred to as xe2x80x9crefractory superalloysxe2x80x9d offer additional strength over the platinum group metals, though at the expense of some oxidation. resistance. These alloys are based on Ir or Rh, with transition metal additions of up to about 20 atomic percent, and are strengthened by a precipitate phase of generic formula M3X, where M is Rh or Ir and X is typically Ti, V, Ta, or Zr, or combinations thereof. Some alloys of this type can withstand 1-2 hour exposures to at least about 1600xc2x0 C. without catastrophic oxidation. Creep life and fatigue life data for these alloys are not readily available currently, but the high strength of these alloys suggests they are superior to some degree over conventional superalloys in both creep life and fatigue life at the temperatures and stress levels relevant to gas turbine airfoil components.
Platinum group metals also have been incorporated into conventional superalloy compositions to produce a class of alloys, herein referred to as xe2x80x9cplatinum-group metal modified superalloysxe2x80x9d, having enhanced oxidation resistance and comparable mechanical properties to conventional superalloys. Typical alloys of this class comprise a conventional superalloy composition to which is added up to about 7 atomic percent of a platinum group metal, such as Ir, Rh, Pt, Pd, and Ru. These alloys comprise the two-phase microstructure typical of conventional superalloys, where a gamma matrix phase (nickel and other dissolved elements, including the platinum group metal elements) is strengthened by precipitates of the gamma prime phase with a general formula of Ni3 (AlTi). Use of materials incorporating platinum-group metals has been limited to date due to the high density and very high cost of these materials in comparison to more conventional airfoil materials.
The selection of a particular alloy for use in a given airfoil design is accomplished based on the critical design requirements for a number of material properties, including strength, toughness, environmental resistance, weight, cost, and others. When one alloy is used to construct the entire airfoil, compromises must be made in the performance of the airfoil because no single alloy possesses ideal values for the long list of properties required for the airfoil application, and because conditions of temperature, stress, impingement of foreign matter, and other factors are not uniform over the entire airfoil surface.
It would be advantageous if the performance of both newly manufactured and repaired airfoils could be improved to better withstand the aggressive stress-temperature combinations present in localized areas on turbine components. However, it would not be desirable if improvements to creep life and fatigue life were effected at the expense of other design critical requirements of the airfoil. For example, a blade made entirely of DSE material would have excellent creep and fatigue properties, but would cost many times the price of a blade made of conventional superalloy material. Therefore, it would be beneficial if turbine airfoils could be improved in a manner that would allow for enhanced performance in regions susceptible to damage due to locally aggressive conditions without significantly detracting from the overall performance of the airfoil.
One embodiment of the present invention provides a method for repairing a gas turbine airfoil, comprising: a. providing an airfoil having specified nominal dimensions, the airfoil comprising a first material having a creep life and a fatigue life, the airfoil further comprising a leading edge section and a trailing edge section; b. removing at least one portion of at least one section of the airfoil to create at least one deficit of material for the airfoil relative to the specified nominal dimensions, the at least one section selected from the group consisting of the leading edge section and the trailing edge section; c. providing at least one insert comprising a second material, the second material having a creep life that is at least substantially equal to the creep life of the first material, and a fatigue life that is at least substantially equal to the fatigue life of the first material; and d. disposing the at least one insert onto the airfoil such that the at least one deficit of material is substantially eliminated.
Another aspect of the invention provides a method for manufacturing a gas turbine airfoil having specified nominal dimensions, a leading edge section, and a trailing edge section, the method comprising: a. providing an airfoil having at least one deficit of material relative to at least one specified nominal dimension, the airfoil comprising a first material having a creep life and a fatigue life, the at least one deficit located in at least one section of the airfoil selected from the group consisting of the leading edge section and the trailing edge section; b. providing at least one insert comprising a second material, the second material having a creep life that is at least substantially equal to the creep life of the first material, and a fatigue life that is at least substantially equal to the fatigue life of the first material; and c. disposing the at least one insert onto the airfoil such that the at least one deficit of material is substantially eliminated.
Another aspect of the invention provides an insert for the manufacture and repair of a gas turbine airfoil, the airfoil comprising a leading edge section and a trailing edge section and having specified nominal dimensions for an airfoil external surface, the insert comprising: an external surface that substantially conforms with the specified nominal dimensions for the airfoil external surface at a section of the airfoil selected from the group consisting of the leading edge section and the trailing edge section; and a material having a creep life at about 1100xc2x0 C. of at least about 50 hours at about 140 MPa and a fatigue life of at least about 4,000 cycles in axial-axial low cycle fatigue testing, the fatigue testing being performed at about 1100xc2x0 C. and with a strain range of about 0.25% at about 20 cycles per minute.
Another aspect of the invention provides a gas turbine airfoil comprising a main body, the main body comprising a first material, the first material having a creep life and a fatigue life, and the airfoil further comprising at least one insert comprising a second material, the second material having a creep life that is at least substantially equal to the creep life of the first material, and a fatigue life that is at least substantially equal to the fatigue life of the first material, the at least one insert joined to the main body, the airfoil comprising a leading edge section and a trailing edge section, wherein the at least one insert comprises at least one portion of at least one section of the airfoil selected from the group consisting of the leading edge section and the trailing edge section.